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Vasculite digital inicial em uma grande coorte multicêntrica de pacientes com lúpus eritematoso sistêmico de início na infância
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
AIAA-97-1195-
Actuation of Trailing Edge Flap in a Wing Model using Piezostack Device* Ramesh Chandra* and Inderjit Chopra** Center for Rotorcraft Education and Research Department of Aerospace Engineering University of Maryland, College Park, MD 20742 Abstract
Actuation of trailing edge flaps of helicopter rotor blades is being pursued these days to alter the blade aerodynamics and hence to control its performance. One key issue is the development of actuators that can efficiently produce sufficient oscillatory flap deflection in rotating and unsteady aerodynamic environments. Spangler and Hall^ demonstrated the feasibility of using a piezoceramic bender element to actuate trailing edge flap to alter the aerodynamics of a helicopter rotor blade. A model to predict the dynamic behavior of the actuator integrated with an airfoil section under non-rotating conditions was developed. The major discrepancy in experimental unsteady pitch moments and theoretical values based on Theordorsen's analysis was shown. For the full-scale implementation, the characteristics of the existing actuators were highlighted as the major limitation. Samak and Chopra^ cantilevered bimorph piezobender element to actuate the trailing edge flap in a Froude scale rotor blade model. The performance of this device was first evaluated in an openjet tunnel and subsequently in a hover condition using a bearingless rotor rig . A flap deflection of 2 degree at 4/rev frequency was achieved at a rotational speed of 900 rpm. To improve upon the performance of this device, Walz and Chopra^ used 4 layer piezo bender elements and a different mechanical design for
This paper presents a methodology to actuate trailing edge flap in a wing model using a high performance piezostack device. A composite wing model was built around this piezostack actuator with integrated mechanical amplification. The free displacement and block force of this high voltage actuator were respectively determined as 28 mil and 16 Ib at 1000 volts. The linear motion of the actuator was converted to rotary motion for flap actuation using a hinge offset mechanism. This model was tested to evaluate the performance of flap actuation without and with aerodynamic forces. The flap deflection was measured using a Hall sensor, and a maximum of 8.5 deg flap deflection was achieved. The influence of free stream velocity and angle of attack on flap deflection was examined. A 30% decrease in flap deflection at free stream velocity at 95 ft/sec was noticed. Angle of attack of 10 deg did not change the flap deflection. Flap deflection was predicted from forcedisplacement relation of actuator and calculated hinge moment and good correlation between experiment and prediction was achieved.
Introduction
+
Copyright © 1997 by Ramesh Chandra and Inderjit Chopra. Presented at the 38th Structures, Structural Dynanics, and Materials Conference and Adaptive Structures Forum, April 7-10,1997, Kissimmee, FL * Currently at Aerostructures, Inc. ,1725 Jeff Davis Highway, Suite 701, Arlington Va 22202. ** Professor and Director 1438
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
conversion of bending motion to rotary motion (to actuate flap ). A flap deflection of 4 degree at 900 rpm was predicted. BenZeev and Chopra^ improved the fabrication of the trailing-edge flap to reduce side frictional force at high rotational speed. However, the hover testing of this model showed a moderate improvement in flap deflection at the operating rotational speed. For these Froude scale rotor models piezobender element, which is basically a moderate displacement and force actuator was used.
mils. Thus the mechanical losses in that device were not too high. That wing model
was tested in an open-jet tunnel and the performance of the flap degraded at high dynamic pressure. Thus, the need to search for better actuators and novel mechanical design to actuate the trailing edge flap in a rotor is obvious. The objective of this research is to develop the trailing edge flap actuation system for a rotor blade using a piezostack actuator in conjunction with integrated mechanical amplification. As a first phase of this research, a fixed wing model with flap actuation is examined .
Piezoceramic, electrostrictive and magnetostrictive devices are high force-low displacement devices and it is possible to trade force for displacement by mechanical means. Samak and Chopra^ explored this possibility of using piezostack actuator with mechanical amplification to actuate leading edge droop of a wing model. The free displacement of the chosen piezostack was 0.75 mil. A mechanical amplification of 25 was used. However, due to mechanical losses , the free displacement of the device was found to be 5 mils as against the calculated value of 18.75 mils at all frequencies. Spencer and Chopra" employed two piezostacks and L-arm amplification to actuate a trailing edge flap in a wing model. Two piezostacks bonded together back to back yielded a free displacement of 1.5 mils which was amplified 10 times. The
Piezostack Actuator Figure 1 shows a schematic of a typical
piezostack device. In this device a large number of piezoelements are bonded together by means of a conducting adhesive.
The displacement of the device in the direction normal to its plane is given as:
v
8 = n d33 —
(i)
where, 8 = displacement , n = number of piezoelements, V = applied voltage and t= thickness of piezoelement and (^33 is piezoelectric constant. It is important to note that for a given piezoceramic material, higher values of transverse displacement can be obtained by using large number of piezoceramic sheets.
calculated value of free displacement was 15 mils as against the experimental value of 11
1439
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Piezoceramic element
Conductive Adhesive Figure 1. Schematic of a typical piezostack actuator
Figure 2 illustrates a sectional view of a
a free displacement of 28 mils and a block force of 16 Ibs of this stack at 1000 volts. The data given by the manufacturer is: free displacement = 700 micron( 27.5 mil ) and block force = 80 N (19.7 Ib). Figure 5 shows the displacement of this actuator at different voltages.
piezostack actuator with integrated mechanical amplification' . This actuator was tested for its force-displacement behavior and Figure 3 shows a schematic of the set-up for this test. Force is applied using dead weights and the displacement is measured using a LVDT (Linear Variable Differential Transformer ). Figure 4 gives force-displacement curve of this stack. Note
L-arm
Amplified displacement
Displacement from the stack
Elastic hinge Figure 2. Piezostack actuator with integrated mechanical amplification
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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Actuator
LVDT
Base suppor t plate
Frame
FApptied force Figure 3. Test set-up for force -displacement behavior of a piezostack actuator 30.0
25.0
Displacement mil
20.0
Displacement 15.0 • 10.0 • 5.0
200
0.0 4
8 12 Force, Ibs
16
20
400
MO
MO
Applied Voltage
Figure 5. Measured free displacement - voltage curve of piezostack actuator
Figure 4. Measured force -displacement curve of piezostack actuator at 1000 volts
this model. This composite wing model consists of a base plate, end ribs, rigid foam core, composite skin and a mechanism (sleeve- push rod ) to convert linear motion to rotary motion. The rigid foam core was made using compression molding technique. The milled core , base plate and sleeves of the mechanism were assembled, two graphite-epoxy prepreg plies were laid on this assembly and cured using metal molds. The process of fabricating the model is described in Figure 7.
Wing Model with Flap Actuation Fabrication A rigid composite wing model of 8 in chord, 16 in span and NACA 0012 airfoil section was designed and built. The flap
actuation was affected using the previously described actuator. The linear motion of the actuator was converted to rotary motion using a push-rod hinge connection with hinge offset. Figure 6 shows a schematic of
1441
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Spring Push rod
Actuator
Base plate
Flap Sleeve
Rod
Figure 6 Schematic of composite wing model with actuated flap Fabrication of rigid foam core by compression molding technique
Machining the rigid foam core to adhesive bond base plate and sleeves
1
Assembly of base plate, end ribs, rigid foam core and sleeves Lay-up of composite plies on the assembly
Fabrication of composite wing model
by matched-die molding technique
Mounting the actuator ,the mechanism to translate linear motion to rotatory motion
and the flap on the wing model
Figure 7 Process of fabricating wing model with flap actuation
In this model, the mechanically amplified stroke of a piezostack actuator is translated to rotary motion to actuate the flap via a push rod. The push rod is supported by two bearings and is free to move in axial direction. It is connected to the flap via a
hinge tube. The offset between the hinge and the connection converts the linear motion to rotary motion. The flap hinge tube is supported by two pins that are inserted in the sleeves, embedded in the model. Figure 8 gives some details of the mechanism.
1442
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Flap
Bearing
Actuator • Figure 8. Details of push rod, hinge tube and flap
Figure 9. Photograph of composite wing model with actuated flap.
Performance of Flap Actuation without Aerodynamic Forces The wing
(Figure 9)
flap deflection (rotation) was measured using a mirror and laser system and Hall sensor. Figure 10 shows the flap deflection at different voltages. The flap deflection of 8
model
with flap actuation was first tested for its
deg. at 1000 volts is experimentally achieved. The calculated value of flap deflection is about 12% higher than the
performance under static conditions by applying a DC voltage to the actuator. The
1443
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
experimental value. The difference between theory and experiment is plausibly due to slackness at the push rod-flap connection.
displacement of the actuator by hinge offset. This flap deflection does not vary with frequency as the linear displacement does not vary with frequency.
10.0 8.0
10.0
•
8.0-
6.0
-y - -0.045793 + 2.353«x fl« 0.99937
Flap
deflection deg
6.0'
4.0 H
Flap
deflection deg
2.0 •
4-0 2.0-
0.0
0
200
400
600
800
1000
0.0 Applied Voltage
0.0
0.5
Figure 10. Flap deflection in composite wing
1.0
1.5
2.0
2.5
3.0
3.5
Sensor output, volt
model at different actuating voltage-static condition
Figure 12 Calibration curve of Hall sensor
o.o ^—Theory 1 • Experiment |
-200.0'
Applied voltage
8.0 '
-400.0-
m
'
25
30
*
Flap
deflection
-800.0-
6.0 4.0 -
-1000.0 0
45
90
135 180 225
270
315
2.0 -
360
Time
Figure 11. Time history of applied voltage to the
10
actuator
15
20
Frequency, Hz
The model was subsequently tested under dynamic conditions. Figure 11 shows time history of maximum applied voltage to the actuator. Note that this stack is configured
Figure 13. Flap deflection in composite wing model at different frequencies
for negative voltage. Flap deflection was measured using a Hall sensor . This sensor is based on Hall effect - the resistance of a coil is influenced by magnetic field and hence consists of a coil and magnet. The coil was mounted on the hinge tube and a cylindrical magnet ( 0.185 in length and 0.118 in diameter ) was inserted in the flap such that the axis of magnet is perpendicular to the plane of the Hall generator (coil). The response of the sensor was monitored using a standard signal conditioner. The drift of the sensor over one hour was within 1%. Figure 12 shows the calibration curve of the Hall sensor. Flap deflection under dynamic condition is given in Figure 13. Note that the theoretical flap deflection was obtained by dividing the experimentally determined linear
Performance of Flap Actuation with Aerodynamic Forces Force-displacement relation of the piezo actuator as given in Figure 4 is written as:
•£l where
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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
u is displacement of actuator F is force corresponding to the displacement u U0 is free displacement (zero force) of actuator Fb is block force (zero displacement)
10.0 8.06.0-
Flap deflection 4.0deg
The linear displacement u is converted to the rotary motion (flap deflection) by the following relation:
2.00.0
1
0
2
Moment, in-lb
(3)
Figure 14. Theoretical flap deflection-moment
where, 8 is flap deflection and e is hinge offset
curve of the actuator
The trailing-edge flap hinge moment is calculated using Glaurt method** as
It is important to note that for a given linear displacement , maximum rotation can be
M a e r o =c h ipV 2 c>
(5)
obtained by minimum hinge offset. However, the force required corresponding to minimum hinge offset will be high. In the present design a hinge offset of 0.178 is used.
where Cf is flap chord and Ch is hinge moment coefficient
Using equations (2) and (3), the flap deflection-moment relation is obtained as:
Hinge moment coefficient depends on flap deflection and C, and is:
chh =
(4)
where,
dc, 0 d8
dc. dc
(6)
{M,Mb} = e{F,Fb}
C, depends on flap deflection and angle of attack and is:
Figure 14 shows the flap deflection moment curve obtained from the forcedisplacement characteristics of a piezostack
C, =-T-r8 +
dc, 5. dc,
d8" ' da
actuator.
a
(7)
For NACA 0012 airfoil, from Ref [8]
*L = 1.337,^ = 0.1 d8 da
(8)
For 20% flap chord, dc
h _ *•= = n0.65, d8
dc
h _= 0.09
dc,
(9)
Using relations (8) and (9),
c h =0.778 + 0.009a
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(10)
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Using relations (4), (5) and (10) 2
0.009
10.0-
——Analysis 1 • Experiment 1
8.0 ____ ^^^^ Flap deflection g.o • ^ deg-pp
(11)
l + 0.77-pV2c* 2 M,
4.0
••*^^^^«
' * iq
Angle of attack=0 deg Frequency=5 Hz
2.0
Note that the flap deflection depends on the force-displacement characteristics of the
20
actuator (free displacement and block force)
Free stream velocity, ft/sec
and the aerodynamic hinge moment ,which in turn depend on free stream velocity and angle of attack.
40
60
80
100
Figure 15 Flap deflection at 5 Hz frequency, 0 deg angle of attack and different velocities
10.0
Testing of Flap Actuation in Wind Tunnel Flap
To validate the analytical model presented
deflection deg-pp
in the previous section for the prediction of
Angle of attack=0 deg Free stream velocity=95 ft/sec
4.0
flap performance, the wing model was tested in an open-jet wind tunnel. The flap performance was evaluated at different angles of attack and free stream velocity .
2.0 ' 0.0
10
Figure 15 shows the dynamic performance of the flap at 5 Hz frequency, zero angle of
15
20
25
Frequency, Hz
attack and different velocities. Note the decrease in the flap deflection with the increase in free stream velocity, the maximum decrease being 30% at 95 ft/sec. Good correlation between analysis and
Figure 16 Rap deflection at 95 ft/sec free stream velocity, 0 deg angle of attack and different frequencies izo
experiment is seen. Figure 16 shows the
mo
influence of frequency on flap performance
at a free stream velocity of 95 ft/sec. The
•Analysis Experiment |
8.0
Flap deflection 6.0 deg-pp
influence of free stream velocity and frequency on flap deflection at angle of attack of 5 and 10 deg is given in figures 1720. It is interesting to note that for this configuration angle of attack does not change the hinge moment significantly and hence the flap deflection is not affected by
4.0 2.0
Angle of attack^s deg Frequency*5 Hz
0.0 20
40
60
80
Free stream velocity, ft/sec
Figure 17 Flap deflection at 5 FIz frequency, 5 deg angle or attack and different velocities
change in angle of attack. This is also obvious by examining equation (10).
1446
30
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
velocity, 10 deg angle of attack and different
frequencies
10.0 8.0-
Flap deflection deg-pp
6.0' 4.0 •
Angle of attack=5 deg Free stream velocity«95 ft/sec
Flap defl. 80 deg
2.0-
6.0 10
15
20
25
30
Frequency, Hz
Figure 18 Flap deflection at 95 ft/sec free stream
100
velocity, 5 deg angle of attack and different
frequencies
——A nalysis
500
under
aerodynamic forces, piezostack actuators with higher free displacement and block force were examined theoretically (Figure 21) . It is interesting to note that an actuator with high free displacement and low block force can create higher flap deflections until certain velocity. It is to be noted that the higher values of free displacement and block force can be realized using available actuators with integrated mechanical amplification of 10. Table 1 shows the force displacement values of such actuators.
8.0-
Angle of attack=10 deg Frequency=5 Hz
2.00.0 20 40 60 80 Free stream velocity, ft/sec
Figure 19 Flap deflection at 5 Hz frequency, 10 deg angle of attack and different velocities
_. 6.0 • Flap deflection deg-pp 4.0
400
For better performance of flaps
I
• Ex xperimcnt |
Flap deflection 6.0 deg-pp 4.0-
300
Figure 21. Flap deflection at different free stream velocities using different piezostack actuators
12.0 10.0 •
200
Free stream velocity, ft/sec
Angle of attack-10 deg
Free stream velocity=90 ft/sec
IS
20
Frequency, Hz
Figure 20 Flap deflection at 90 ft/sec free stream Table 1. Details of piezostack actuators
Actuator P287.70 P290.00* P246.70* P245.70* P844.60* Dual Stack*
Free displ. mil 28 39 47 47 35 15
Block Force Ib 16 11 275 44 66 130
Diameter in 1.57 0.71 0.78
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Length in 5.57 5.00 5.39
Cost $ 2265
3190 4940 2250 4565
Voltage high high high high
low lOw
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
* These actuators are not available with integrated mechanical amplification. However, a mechanical amplification of 10 is used for these calculations 4. Ben-Zeev, O., and Chopra, I.," Advances in the Development of an Intelligent Helicopter Rotor Employing Trailing-Edge Flaps" Smart Structures and Materials, Vol. 5. No. 1, February 1996, pp. 11-25.
Conclusions A piezostack actuator with integrated mechanical amplification was used to build composite wing model with flap. The model was tested for flap performance without and with aerodynamic forces. Flap deflections of approximately 8 deg were obtained in the absence of aerodynamic forces . Aerodynamic hinge moments at 0 deg angle of attack and free stream velocity of 100 ft/sec reduced the mean value of flap deflection by about 30% . It appears feasible to build a full-scale traling-edge flap with specially designed piezostacks for real life conditions.
5. Samak, D. K., and Chopra, I., " Design of High Force, High Displacement Actuators for Helicopter Rotors", Smart Structures and Materials, Vol. 5, No. 1, February 1996, pp. 58-67. 6. Spencer, B., and Chopra, I., " Design and Testing of a Smart Trailing Edge Flap using Piezoelectric Stacks", Proceedings of SPIE Conference on Smart Structures and Materials, Feb. 25-29 , 1996, San Diego, California.
Acknowledgement
7. Products for Micropositioning, Physik Instrumente
This work was sponsored by US Army Aeroflightdynamics Directorate at Ames under Grant No NCC 25160: Technical Monitor Dr. Chee Tung
8. Abbott, I. H., and Von Doenhoff, A. E., " of Wing Sections", Dover Publications, Inc. New York. Theory
References 1. Spangler, Jr., R. L. and Hall, S. R., "Piezoelectric Actuators for Helicopter Rotor Control", Proceedings of the 30th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, AIAA, Washington D. C. April 1990.
2. Samak, D. K., and Chopra, L, " A Feasibility Study to Build a Smart Rotor: Trailing Edge Flap Actuation" SPIE Conference on Smart Structures and Materials, Feb. 1-4 ,1993, Albuquerque, New Mexico. 3. Walz, C. , and Chopra, I. ," Design and Testing of a Helicopter Rotor Model with
Smart Trailing Edge Flaps", Proceedings of the 35th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, AIAA, Washington D. C. April 1995. 1448
AIAA-97-1195-
Actuation of Trailing Edge Flap in a Wing Model using Piezostack Device* Ramesh Chandra* and Inderjit Chopra** Center for Rotorcraft Education and Research Department of Aerospace Engineering University of Maryland, College Park, MD 20742 Abstract
Actuation of trailing edge flaps of helicopter rotor blades is being pursued these days to alter the blade aerodynamics and hence to control its performance. One key issue is the development of actuators that can efficiently produce sufficient oscillatory flap deflection in rotating and unsteady aerodynamic environments. Spangler and Hall^ demonstrated the feasibility of using a piezoceramic bender element to actuate trailing edge flap to alter the aerodynamics of a helicopter rotor blade. A model to predict the dynamic behavior of the actuator integrated with an airfoil section under non-rotating conditions was developed. The major discrepancy in experimental unsteady pitch moments and theoretical values based on Theordorsen's analysis was shown. For the full-scale implementation, the characteristics of the existing actuators were highlighted as the major limitation. Samak and Chopra^ cantilevered bimorph piezobender element to actuate the trailing edge flap in a Froude scale rotor blade model. The performance of this device was first evaluated in an openjet tunnel and subsequently in a hover condition using a bearingless rotor rig . A flap deflection of 2 degree at 4/rev frequency was achieved at a rotational speed of 900 rpm. To improve upon the performance of this device, Walz and Chopra^ used 4 layer piezo bender elements and a different mechanical design for
This paper presents a methodology to actuate trailing edge flap in a wing model using a high performance piezostack device. A composite wing model was built around this piezostack actuator with integrated mechanical amplification. The free displacement and block force of this high voltage actuator were respectively determined as 28 mil and 16 Ib at 1000 volts. The linear motion of the actuator was converted to rotary motion for flap actuation using a hinge offset mechanism. This model was tested to evaluate the performance of flap actuation without and with aerodynamic forces. The flap deflection was measured using a Hall sensor, and a maximum of 8.5 deg flap deflection was achieved. The influence of free stream velocity and angle of attack on flap deflection was examined. A 30% decrease in flap deflection at free stream velocity at 95 ft/sec was noticed. Angle of attack of 10 deg did not change the flap deflection. Flap deflection was predicted from forcedisplacement relation of actuator and calculated hinge moment and good correlation between experiment and prediction was achieved.
Introduction
+
Copyright © 1997 by Ramesh Chandra and Inderjit Chopra. Presented at the 38th Structures, Structural Dynanics, and Materials Conference and Adaptive Structures Forum, April 7-10,1997, Kissimmee, FL * Currently at Aerostructures, Inc. ,1725 Jeff Davis Highway, Suite 701, Arlington Va 22202. ** Professor and Director 1438
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
conversion of bending motion to rotary motion (to actuate flap ). A flap deflection of 4 degree at 900 rpm was predicted. BenZeev and Chopra^ improved the fabrication of the trailing-edge flap to reduce side frictional force at high rotational speed. However, the hover testing of this model showed a moderate improvement in flap deflection at the operating rotational speed. For these Froude scale rotor models piezobender element, which is basically a moderate displacement and force actuator was used.
mils. Thus the mechanical losses in that device were not too high. That wing model
was tested in an open-jet tunnel and the performance of the flap degraded at high dynamic pressure. Thus, the need to search for better actuators and novel mechanical design to actuate the trailing edge flap in a rotor is obvious. The objective of this research is to develop the trailing edge flap actuation system for a rotor blade using a piezostack actuator in conjunction with integrated mechanical amplification. As a first phase of this research, a fixed wing model with flap actuation is examined .
Piezoceramic, electrostrictive and magnetostrictive devices are high force-low displacement devices and it is possible to trade force for displacement by mechanical means. Samak and Chopra^ explored this possibility of using piezostack actuator with mechanical amplification to actuate leading edge droop of a wing model. The free displacement of the chosen piezostack was 0.75 mil. A mechanical amplification of 25 was used. However, due to mechanical losses , the free displacement of the device was found to be 5 mils as against the calculated value of 18.75 mils at all frequencies. Spencer and Chopra" employed two piezostacks and L-arm amplification to actuate a trailing edge flap in a wing model. Two piezostacks bonded together back to back yielded a free displacement of 1.5 mils which was amplified 10 times. The
Piezostack Actuator Figure 1 shows a schematic of a typical
piezostack device. In this device a large number of piezoelements are bonded together by means of a conducting adhesive.
The displacement of the device in the direction normal to its plane is given as:
v
8 = n d33 —
(i)
where, 8 = displacement , n = number of piezoelements, V = applied voltage and t= thickness of piezoelement and (^33 is piezoelectric constant. It is important to note that for a given piezoceramic material, higher values of transverse displacement can be obtained by using large number of piezoceramic sheets.
calculated value of free displacement was 15 mils as against the experimental value of 11
1439
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Piezoceramic element
Conductive Adhesive Figure 1. Schematic of a typical piezostack actuator
Figure 2 illustrates a sectional view of a
a free displacement of 28 mils and a block force of 16 Ibs of this stack at 1000 volts. The data given by the manufacturer is: free displacement = 700 micron( 27.5 mil ) and block force = 80 N (19.7 Ib). Figure 5 shows the displacement of this actuator at different voltages.
piezostack actuator with integrated mechanical amplification' . This actuator was tested for its force-displacement behavior and Figure 3 shows a schematic of the set-up for this test. Force is applied using dead weights and the displacement is measured using a LVDT (Linear Variable Differential Transformer ). Figure 4 gives force-displacement curve of this stack. Note
L-arm
Amplified displacement
Displacement from the stack
Elastic hinge Figure 2. Piezostack actuator with integrated mechanical amplification
1440
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Actuator
LVDT
Base suppor t plate
Frame
FApptied force Figure 3. Test set-up for force -displacement behavior of a piezostack actuator 30.0
25.0
Displacement mil
20.0
Displacement 15.0 • 10.0 • 5.0
200
0.0 4
8 12 Force, Ibs
16
20
400
MO
MO
Applied Voltage
Figure 5. Measured free displacement - voltage curve of piezostack actuator
Figure 4. Measured force -displacement curve of piezostack actuator at 1000 volts
this model. This composite wing model consists of a base plate, end ribs, rigid foam core, composite skin and a mechanism (sleeve- push rod ) to convert linear motion to rotary motion. The rigid foam core was made using compression molding technique. The milled core , base plate and sleeves of the mechanism were assembled, two graphite-epoxy prepreg plies were laid on this assembly and cured using metal molds. The process of fabricating the model is described in Figure 7.
Wing Model with Flap Actuation Fabrication A rigid composite wing model of 8 in chord, 16 in span and NACA 0012 airfoil section was designed and built. The flap
actuation was affected using the previously described actuator. The linear motion of the actuator was converted to rotary motion using a push-rod hinge connection with hinge offset. Figure 6 shows a schematic of
1441
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Spring Push rod
Actuator
Base plate
Flap Sleeve
Rod
Figure 6 Schematic of composite wing model with actuated flap Fabrication of rigid foam core by compression molding technique
Machining the rigid foam core to adhesive bond base plate and sleeves
1
Assembly of base plate, end ribs, rigid foam core and sleeves Lay-up of composite plies on the assembly
Fabrication of composite wing model
by matched-die molding technique
Mounting the actuator ,the mechanism to translate linear motion to rotatory motion
and the flap on the wing model
Figure 7 Process of fabricating wing model with flap actuation
In this model, the mechanically amplified stroke of a piezostack actuator is translated to rotary motion to actuate the flap via a push rod. The push rod is supported by two bearings and is free to move in axial direction. It is connected to the flap via a
hinge tube. The offset between the hinge and the connection converts the linear motion to rotary motion. The flap hinge tube is supported by two pins that are inserted in the sleeves, embedded in the model. Figure 8 gives some details of the mechanism.
1442
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Flap
Bearing
Actuator • Figure 8. Details of push rod, hinge tube and flap
Figure 9. Photograph of composite wing model with actuated flap.
Performance of Flap Actuation without Aerodynamic Forces The wing
(Figure 9)
flap deflection (rotation) was measured using a mirror and laser system and Hall sensor. Figure 10 shows the flap deflection at different voltages. The flap deflection of 8
model
with flap actuation was first tested for its
deg. at 1000 volts is experimentally achieved. The calculated value of flap deflection is about 12% higher than the
performance under static conditions by applying a DC voltage to the actuator. The
1443
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
experimental value. The difference between theory and experiment is plausibly due to slackness at the push rod-flap connection.
displacement of the actuator by hinge offset. This flap deflection does not vary with frequency as the linear displacement does not vary with frequency.
10.0 8.0
10.0
•
8.0-
6.0
-y - -0.045793 + 2.353«x fl« 0.99937
Flap
deflection deg
6.0'
4.0 H
Flap
deflection deg
2.0 •
4-0 2.0-
0.0
0
200
400
600
800
1000
0.0 Applied Voltage
0.0
0.5
Figure 10. Flap deflection in composite wing
1.0
1.5
2.0
2.5
3.0
3.5
Sensor output, volt
model at different actuating voltage-static condition
Figure 12 Calibration curve of Hall sensor
o.o ^—Theory 1 • Experiment |
-200.0'
Applied voltage
8.0 '
-400.0-
m
'
25
30
*
Flap
deflection
-800.0-
6.0 4.0 -
-1000.0 0
45
90
135 180 225
270
315
2.0 -
360
Time
Figure 11. Time history of applied voltage to the
10
actuator
15
20
Frequency, Hz
The model was subsequently tested under dynamic conditions. Figure 11 shows time history of maximum applied voltage to the actuator. Note that this stack is configured
Figure 13. Flap deflection in composite wing model at different frequencies
for negative voltage. Flap deflection was measured using a Hall sensor . This sensor is based on Hall effect - the resistance of a coil is influenced by magnetic field and hence consists of a coil and magnet. The coil was mounted on the hinge tube and a cylindrical magnet ( 0.185 in length and 0.118 in diameter ) was inserted in the flap such that the axis of magnet is perpendicular to the plane of the Hall generator (coil). The response of the sensor was monitored using a standard signal conditioner. The drift of the sensor over one hour was within 1%. Figure 12 shows the calibration curve of the Hall sensor. Flap deflection under dynamic condition is given in Figure 13. Note that the theoretical flap deflection was obtained by dividing the experimentally determined linear
Performance of Flap Actuation with Aerodynamic Forces Force-displacement relation of the piezo actuator as given in Figure 4 is written as:
•£l where
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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
u is displacement of actuator F is force corresponding to the displacement u U0 is free displacement (zero force) of actuator Fb is block force (zero displacement)
10.0 8.06.0-
Flap deflection 4.0deg
The linear displacement u is converted to the rotary motion (flap deflection) by the following relation:
2.00.0
1
0
2
Moment, in-lb
(3)
Figure 14. Theoretical flap deflection-moment
where, 8 is flap deflection and e is hinge offset
curve of the actuator
The trailing-edge flap hinge moment is calculated using Glaurt method** as
It is important to note that for a given linear displacement , maximum rotation can be
M a e r o =c h ipV 2 c>
(5)
obtained by minimum hinge offset. However, the force required corresponding to minimum hinge offset will be high. In the present design a hinge offset of 0.178 is used.
where Cf is flap chord and Ch is hinge moment coefficient
Using equations (2) and (3), the flap deflection-moment relation is obtained as:
Hinge moment coefficient depends on flap deflection and C, and is:
chh =
(4)
where,
dc, 0 d8
dc. dc
(6)
{M,Mb} = e{F,Fb}
C, depends on flap deflection and angle of attack and is:
Figure 14 shows the flap deflection moment curve obtained from the forcedisplacement characteristics of a piezostack
C, =-T-r8 +
dc, 5. dc,
d8" ' da
actuator.
a
(7)
For NACA 0012 airfoil, from Ref [8]
*L = 1.337,^ = 0.1 d8 da
(8)
For 20% flap chord, dc
h _ *•= = n0.65, d8
dc
h _= 0.09
dc,
(9)
Using relations (8) and (9),
c h =0.778 + 0.009a
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(10)
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
Using relations (4), (5) and (10) 2
0.009
10.0-
——Analysis 1 • Experiment 1
8.0 ____ ^^^^ Flap deflection g.o • ^ deg-pp
(11)
l + 0.77-pV2c* 2 M,
4.0
••*^^^^«
' * iq
Angle of attack=0 deg Frequency=5 Hz
2.0
Note that the flap deflection depends on the force-displacement characteristics of the
20
actuator (free displacement and block force)
Free stream velocity, ft/sec
and the aerodynamic hinge moment ,which in turn depend on free stream velocity and angle of attack.
40
60
80
100
Figure 15 Flap deflection at 5 Hz frequency, 0 deg angle of attack and different velocities
10.0
Testing of Flap Actuation in Wind Tunnel Flap
To validate the analytical model presented
deflection deg-pp
in the previous section for the prediction of
Angle of attack=0 deg Free stream velocity=95 ft/sec
4.0
flap performance, the wing model was tested in an open-jet wind tunnel. The flap performance was evaluated at different angles of attack and free stream velocity .
2.0 ' 0.0
10
Figure 15 shows the dynamic performance of the flap at 5 Hz frequency, zero angle of
15
20
25
Frequency, Hz
attack and different velocities. Note the decrease in the flap deflection with the increase in free stream velocity, the maximum decrease being 30% at 95 ft/sec. Good correlation between analysis and
Figure 16 Rap deflection at 95 ft/sec free stream velocity, 0 deg angle of attack and different frequencies izo
experiment is seen. Figure 16 shows the
mo
influence of frequency on flap performance
at a free stream velocity of 95 ft/sec. The
•Analysis Experiment |
8.0
Flap deflection 6.0 deg-pp
influence of free stream velocity and frequency on flap deflection at angle of attack of 5 and 10 deg is given in figures 1720. It is interesting to note that for this configuration angle of attack does not change the hinge moment significantly and hence the flap deflection is not affected by
4.0 2.0
Angle of attack^s deg Frequency*5 Hz
0.0 20
40
60
80
Free stream velocity, ft/sec
Figure 17 Flap deflection at 5 FIz frequency, 5 deg angle or attack and different velocities
change in angle of attack. This is also obvious by examining equation (10).
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30
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
velocity, 10 deg angle of attack and different
frequencies
10.0 8.0-
Flap deflection deg-pp
6.0' 4.0 •
Angle of attack=5 deg Free stream velocity«95 ft/sec
Flap defl. 80 deg
2.0-
6.0 10
15
20
25
30
Frequency, Hz
Figure 18 Flap deflection at 95 ft/sec free stream
100
velocity, 5 deg angle of attack and different
frequencies
——A nalysis
500
under
aerodynamic forces, piezostack actuators with higher free displacement and block force were examined theoretically (Figure 21) . It is interesting to note that an actuator with high free displacement and low block force can create higher flap deflections until certain velocity. It is to be noted that the higher values of free displacement and block force can be realized using available actuators with integrated mechanical amplification of 10. Table 1 shows the force displacement values of such actuators.
8.0-
Angle of attack=10 deg Frequency=5 Hz
2.00.0 20 40 60 80 Free stream velocity, ft/sec
Figure 19 Flap deflection at 5 Hz frequency, 10 deg angle of attack and different velocities
_. 6.0 • Flap deflection deg-pp 4.0
400
For better performance of flaps
I
• Ex xperimcnt |
Flap deflection 6.0 deg-pp 4.0-
300
Figure 21. Flap deflection at different free stream velocities using different piezostack actuators
12.0 10.0 •
200
Free stream velocity, ft/sec
Angle of attack-10 deg
Free stream velocity=90 ft/sec
IS
20
Frequency, Hz
Figure 20 Flap deflection at 90 ft/sec free stream Table 1. Details of piezostack actuators
Actuator P287.70 P290.00* P246.70* P245.70* P844.60* Dual Stack*
Free displ. mil 28 39 47 47 35 15
Block Force Ib 16 11 275 44 66 130
Diameter in 1.57 0.71 0.78
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Length in 5.57 5.00 5.39
Cost $ 2265
3190 4940 2250 4565
Voltage high high high high
low lOw
Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.
* These actuators are not available with integrated mechanical amplification. However, a mechanical amplification of 10 is used for these calculations 4. Ben-Zeev, O., and Chopra, I.," Advances in the Development of an Intelligent Helicopter Rotor Employing Trailing-Edge Flaps" Smart Structures and Materials, Vol. 5. No. 1, February 1996, pp. 11-25.
Conclusions A piezostack actuator with integrated mechanical amplification was used to build composite wing model with flap. The model was tested for flap performance without and with aerodynamic forces. Flap deflections of approximately 8 deg were obtained in the absence of aerodynamic forces . Aerodynamic hinge moments at 0 deg angle of attack and free stream velocity of 100 ft/sec reduced the mean value of flap deflection by about 30% . It appears feasible to build a full-scale traling-edge flap with specially designed piezostacks for real life conditions.
5. Samak, D. K., and Chopra, I., " Design of High Force, High Displacement Actuators for Helicopter Rotors", Smart Structures and Materials, Vol. 5, No. 1, February 1996, pp. 58-67. 6. Spencer, B., and Chopra, I., " Design and Testing of a Smart Trailing Edge Flap using Piezoelectric Stacks", Proceedings of SPIE Conference on Smart Structures and Materials, Feb. 25-29 , 1996, San Diego, California.
Acknowledgement
7. Products for Micropositioning, Physik Instrumente
This work was sponsored by US Army Aeroflightdynamics Directorate at Ames under Grant No NCC 25160: Technical Monitor Dr. Chee Tung
8. Abbott, I. H., and Von Doenhoff, A. E., " of Wing Sections", Dover Publications, Inc. New York. Theory
References 1. Spangler, Jr., R. L. and Hall, S. R., "Piezoelectric Actuators for Helicopter Rotor Control", Proceedings of the 30th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, AIAA, Washington D. C. April 1990.
2. Samak, D. K., and Chopra, L, " A Feasibility Study to Build a Smart Rotor: Trailing Edge Flap Actuation" SPIE Conference on Smart Structures and Materials, Feb. 1-4 ,1993, Albuquerque, New Mexico. 3. Walz, C. , and Chopra, I. ," Design and Testing of a Helicopter Rotor Model with
Smart Trailing Edge Flaps", Proceedings of the 35th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, AIAA, Washington D. C. April 1995. 1448